Gas turbine engines can propel aircraft at supersonic speeds. However, the gas turbine engines generally operate on subsonic air flow in the range of about Mach 0.3 to 0.6 at the upstream face of an engine. The inlet of the engine functions to decelerate the incoming airflow to a speed compatible with the requirements of the gas turbine engine. In order to do this, the inlet has a compression surface and a corresponding flow path, used to decelerate the supersonic flow into a strong terminal shock. A diffuser further decelerates the resulting flow from the strong terminal shock to a speed corresponding to the requirements of the gas turbine engine.
A measurement of inlet operation efficiency is the total pressure lost in the air stream between the entrance side and the discharge side of the inlet. The total pressure recovery of an inlet is defined by a ratio of the total pressure at the discharge to the total pressure at the free stream. Maximizing the total pressure recovery leads to maximizing gross engine thrust, thus improving the performance of the propulsion system. Traditional inlet design methods have aimed at maximizing total pressure recovery. This traditional approach, however, often results in a complex inlet design with high drag.
A traditional approach to supersonic inlet design typically employs shock-on-lip focusing. As understood by those of skill in the art, shock-on-lip focusing involves designing a compression surface configuration of an external compression inlet such that the inlet-generated shocks (that occur at a supersonic design cruise speed) meet at a location immediately forward of the cowl highlight or the cowl lip. The advantages of shock-on-lip focusing include better pressure recovery and low flow spillage drag.
Also, when using shock-on-lip focusing, the cowl lip angle of the cowling may be aligned with the local supersonic flow in the vicinity of the terminal shock in order to prevent formation of an adverse subsonic diffuser flow area profile or a complex internal shock structure in the lip region. If this is not done, a complex internal shock structure and an adverse subsonic diffuser flow area profile may result, possibly reducing the inlet pressure recovery and flow pumping efficiency, as well as undermining diffuser flow stability.
As understood in the art, as supersonic design speed increases, so will the amount of compression necessary to decelerate the flow to a fixed terminal shock Mach number. Additional compression requires more flow-turning off of the inlet axis, resulting in a corresponding increase in the cowl lip angle (in order to align the cowl lip angle with the local flow at the terminal shock). FIG. 1 schematically illustrates a side view of a conventional inlet 1. Inlet 1 has a compression surface 10 and a cowl lip 11. Cowl lip 11 is positioned such that both an initial shock and a terminal shock from compression surface 10 meet at a point before the cowl lip 11. A cowl lip angle 12 is formed when the cowl lip 11 is aligned with the local flow. As mentioned, when the supersonic design speed increases, the amount of compression needed to decelerate the flow to a fixed terminal shock Mach number also increases, resulting in an increase in cowl lip angle. Any increase in cowl lip angle results in additional inlet frontal area, which increases inlet drag as speed increases. This adverse trend is a key reason why conventional external compression inlets lose viability at high supersonic Mach numbers.
One way to control lip drag, as discussed in U.S. Pat. No. 6,793,175 to Sanders, involves configuring the inlet to minimize the shape and size of the cowl. The configuration of the inlet initially resembles a circumferential sector of an axisymmetric intake, but switches the location of compression surface to the outer radius and disposes the cowling on the inner radius in a higher performance, 3-D geometry. The fact that the cowl is located on the inner radius reduces the physical arc of the cowl. Problems with this method include the aircraft integration challenges created by the 3-D geometry, such as the fact that the cross-sectional shape may be more difficult to integrate from a packaging perspective compared to an equivalent axisymmetric design for podded propulsion systems. In addition, the complex inlet shape is likely to create complex distortion patterns that require either large scale mitigating techniques in the subsonic diffuser or the use of engines with more robust operability characteristics.
Another way to control drag by reducing the cowl lip angle is based on decreasing the flow turn angle by increasing the inlet terminal shock Mach number. The improvement in drag reduction is often negated by the reduction in pressure recovery resulting from the stronger terminal shock. In addition, increasing the terminal shock Mach number at the base of the shock also encounters significant limitations in practice due to viscous flow effects. Higher terminal shock Mach numbers at the base of the shock aggravate the shock-boundary layer interaction and reduce shock base boundary layer health. The increase in shock strength in the base region also reduces inlet buzz margin, reducing subcritical flow throttling capability. Additionally, the increase in terminal shock Mach number will most likely require a complex boundary layer management or inlet control system.
Inlet compression surfaces are typically grouped into two types: straight or isentropic. A straight surface has a flat ramp or conic sections that produce discrete oblique or conic shocks, while an isentropic surface has a continuously curved surface that produces a continuum of infinitesimally weak shocklets during the compression process. Theoretically, a traditional isentropic compression surface can have better pressure recovery than a straight surface designed to the same operating conditions, but real viscous effects can reduce the overall performance of the isentropic surface inlets and result in poorer boundary layer health.